1. Field of the Invention
The present invention relates generally to attitude and heading reference detecting apparatus and, more particularly, is directed to an attitude and heading reference detecting apparatus having a gyro, an accelerometer and a magnetic azimuth sensor for use with navigation vehicles such as an aircraft, an automobile or the like.
2. Description of the Prior Art
In the art, an aircraft is provided with a DG (directional gyro) and a magnetic compass (or flux valve compass), each of which is used to indicate the azimuth of the aircraft, a VG (vertical gyro) for indicating the attitude and heading reference angle of the aircraft and a turning meter for indicating the turning angular velocity or speed and a bank angle around the vertical axis of the aircraft, whereby the flight sense of the pilot can be supplemented, thus making it possible to secure safety of navigation of aircraft under any flight conditions.
Initially, the conventional attitude and heading reference detecting apparatus will be described with reference to FIGS. 1 to 4.
FIG. 1 is a perspective view illustrating an example of a directional gyro used in the today's aircraft. As illustrated in FIG. 1, this conventional directional gyro comprises a gyro rotor 101 having spin shafts (axes) 100 and 100a arranged in the substantially horizontal direction which is rotated at high speed and a horizontal ring 102 which is a gyro case for supporting the gyro rotor 101 at its spin shafts 100 and 100a to be freely rotatable. The horizontal ring 102 has horizontal shafts (axes) 103 and 103a disposed at the positions perpendicular to the spin shafts 100 and 100a. These horizontal shafts 103 and 103a are rotatably engaged with horizontal shaft bearings 105 and 105a (horizontal shaft bearing 105a is not shown) fixed to a vertical ring 104 at the positions corresponding to the horizontal shafts 103 and 103a, respectively. The vertical ring 104 is provided with vertical shafts (axes) 106 and 106a extended in the up and down direction at the positions perpendicular to the horizontal shaft bearings 105 and 105a. These vertical shafts 106 and 106a are rotatably engaged with vertical shaft bearings 107 and 107a (vertical shaft bearing 107a is not shown) fixed to a base plate 107B that is fixed to the aircraft at the positions corresponding to the vertical shafts 106 and 106a. To the upper vertical shaft 106a, there are mounted an erecting torquer 108 and a compass card 109, while a synchro receiver 110 and a synchro transmitter 111 are mounted to the lower vertical shaft 106.
The horizontal ring 102 is provided with an electrolytic level 112 which detects the inclination of the spin shafts 100 and 100a relative to the horizontal plane. The output from the electrolytic level 112 is fed through an amplifier 113 back to the erecting torquer 108, thus always maintaining the spin shafts 100 and 100a of the gyro rotor 101 in the horizontal plane. This loop is referred to as an erection loop. A flux valve 114 is adapted to magnetically detect the azimuth angle of the body of the aircraft and produces the magnetic azimuth output. This magnetic azimuth output therefrom is supplied to the synchro receiver 110 which produces a deviation signal between the magnetic azimuth output and the azimuth of the spin shafts 100 and 100a, that is, the gyro azimuth. This deviation signal is fed through an amplifier 114A back to a slaving torquer 115 provided at the horizontal shaft 103a, thus causing the gyro azimuth to coincide with the magnetic azimuth. This loop is referred to as an azimuth slaving loop. Under severe movement of the aircraft, the gyro azimuth is generated and an azimuth angle error caused by the gyro drift is restricted or slaved to the magnetic azimuth derived from the flux valve 114, thus maintaining the accuracy of the azimuth of the aircraft's body. The azimuth of the aircraft's body is read through the compass card 109 mounted to the vertical shaft 106a.
FIG. 2 illustrates an example of a conventional gyroscopic horizon (vertical gyro) for detecting an inclination angle (an angle of roll and an angle of pitch) of the body of an aircraft. In this example, an inner gimbal 132 includes therein a gyro rotor 130 having a spin shaft 131 in the substantially vertical direction and which is rotated at high speed. The inner gimbal 132 is provided at its horizontal positions perpendicular to the spin shaft 131 with pitch shafts 133 and 133a. These pitch shafts 133 and 133a are rotatably engaged with pitch shaft bearings 134 and 134a (pitch shaft bearing 134 is not shown) fixed to an outer gimbal 135 at the positions corresponding to the pitch shafts 134 and 134a. The outer gimbal 135 is provided at its positions perpendicular to the above-mentioned pitch shafts 133 and 133a with roll shafts 136 and 136a. These roll shafts 136 and 136a are rotatably engaged with roll shaft bearings 137 and 137a fixed to roll base tables 138 and 138a attached to the body of aircraft in the direction from its nose to the tail. The inner gimbal 132 includes a roll electrolytic level 139 for detecting the inclination of the spin shaft 131 relative to the horizontal plane around the roll shafts 136 and 136a and a pitch electrolytic level 142 for detecting the inclination of the spin shaft 131 around the pitch shafts 133 and 133a.
The output from the roll electrolytic level 139 is fed through a roll amplifier 140 back to a roll torquer 141 attached to the pitch shaft 133 so as to make the output from the roll electrolytic level 139 zero. This loop is referred to as a roll erecting system. On the other hand, the output from the pitch electrolytic level 142 is fed through a pitch amplifier 143 back to a pitch torquer 144 attached to the roll shaft 136, thus maintaining the inclination of the spin shaft 131 around the pitch shafts 133 and 133a to be zero. This loop is referred to as a pitch erecting system. The roll angle of the body of the aircraft is delivered from a roll angle transmitter 145 mounted to the roll shaft 136a, while the pitch angle is delivered from a pitch angle transmitter 146 mounted to the pitch shaft 133a.
FIG. 3 illustrates an indication section of a conventional turning meter of the aircraft. A lubber line 151 and a needle 152 of a turning angle speed indication portion 153 are used to indicate the turning speed of the aircraft measured by a gyro shown in FIG. 4. The lower half of the indication section is used as a bank angle indication section 154 which indicates a bank angle on the basis of the position of a ball 155 sealed into an annular portion having a curvature.
FIG. 4 illustrates a rate gyro which constructs the above-mentioned turning meter for detecting the turning angle speed of the body of the aircraft. Referring to FIG. 4, a gyro case 171 housing therein a gyro rotor 170 is provided at its positions perpendicular to the axis XX' of a spin shaft 172 of the gyro rotor 177 with output shafts 173 and 173a. These output shafts 173 and 173a are rotatably engaged with output shaft bearings 175 and 175a fixed to a base table 174 which is fixed to the body of the aircraft. A restoring spring 176 and a damping pot 177 are provided between the gyro case 171 and the base table 174.
If a turning angular speed .OMEGA. is applied around an input axis ZZ' perpendicular to both an output axis YY' and the axis XX' of the spin shaft 172, the gyro action causes a torque proportional to the turning angular speed .OMEGA. to be produced around the output axis YY'. This torque causes the gyro case 171 with its association elements to be rotated around the output axis YY' so that the restoring spring 176 produces a torque in response to the displacement angle of the gyro case 171, thus producing a balanced state. In other words, the input angular speed .OMEGA. is transformed into a rotation angle around the output axis YY' and this displacement angle is indicated by a needle 178 (corresponding to the needle 152 shown in FIG. 3) attached to the output shaft 173a.
The conventional apparatus as mentioned above, however, is complicated in mechanism, takes a lot of time and skill for its assembly and adjustment and is also high in cost. Further, since this conventional apparatus requires parts to be abraded such as a ball bearing, sliding electrical contacts and so on, this apparatus must undergo the routine maintenance and inspection and is also affected easily by vibration and shock.
FIG. 5 is a partially cross-sectional view illustrating an example of a magnetic azimuth sensor 7 which is what might be called flux valve compass. In this example, a magnetic azimuth detecting section 7-3 is suspended from the center of a base table 7-1 through a universal joint 7-2 just like a pendulum. A bowl portion 7-4 is attached to the lower surface of the base table 7-1 to form a case or container which houses therein the magnetic azimuth detecting section 7-3. This case is filled with damping oil 7-5, whereby pendulum movement of the magnetic azimuth detecting section 7-3 is attenuated. In spite of the inclination of the navigation vehicle, the universal joint 7-2 acts to always maintain the magnetic azimuth detecting section 7-3 in the horizontal plane, thus preventing the azimuth error from being caused by the inclination of the navigation vehicle.
FIG. 6 is a perspective view of the magnetic azimuth detecting section 7-3 shown in FIG. 5. As shown in FIG. 6, this magnetic azimuth detecting section 7-3 is provided at its central portion with an exciter coil 70 which is formed of windings to be excited by an A.C. voltage source (not shown). Three pick-up coils 72-1, 72-2 and 72-3 are formed by windings wound around three spokes 71-1, 71-2 and 71-3 made of material having high magnetic permeability and fixed to the exciter coil 70 with an equal angular distance between adjacent ones and produce A.C. voltages corresponding to the azimuth of the navigation vehicle with a phase of 120.degree..
However, since this conventional magnetic azimuth sensor 7 has the magnetic azimuth detecting section 7-3 supported by the universal point 7-2 and requires the damping oil 7-5 for damping the magnetic azimuth detecting section 7-3 as mentioned above, it becomes large in size, heavy in weight and is also made high in cost.